1. Field
The disclosed embodiments relate to the field of the assembly of an aircraft fuselage. More particularly, the disclosed embodiments relate to the assembly of panels of an aircraft fuselage when the forces to be transmitted between the fuselage panels are high.
2. Brief Description of Related Developments
An aircraft fuselage is a structure usually comprising substantially cylindrical sections butted against one another along join lines, called circumferential joins, defining planes perpendicular to the longitudinal axis of the fuselage, the sections themselves usually each consisting of several panels also assembled together along join lines, called longitudinal joins, oriented substantially along generatrixes of said fuselage.
These two types of joins are zones of weakness of the fuselage that must be made to withstand the heavy stresses to which the fuselage is subjected in flight.
Usually, the longitudinal joins of the fuselage are made by overlapping the ends of the two panels to be assembled, because the extra thickness linked to the outer panel, being in the direction of aerodynamic flow of the fuselage, does not generate any penalty on the performance of the aircraft.
On the other hand, at a circumferential join, an overlap of two panels would generate an unacceptable aerodynamic discontinuity, which justifies end-to-end assembly.
Therefore, to assemble two panels 2a, 2b, together end-to-end, it is known practice in the prior art, as illustrated in FIG. 1, to add, inside the fuselage and at a circumferential join 8, a reinforcing element, called a ring 3, having the shape of a plate partially overlapping the ends 22a, 22b of the two panels 2a, 2b situated facing one another at the join 8, in order to allow the transmission of the forces to be continuous between said two panels. One face 31 of said ring rests on inner faces 21a, 21b of the panels 2a, 2b, and the ring 3 is assembled and fixed to said two panels by means of fasteners 7, such as rivets, preferably with a flat head to maintain a surface state taking account of the aerodynamic stresses of the aircraft.
To increase the strength of the fuselage, while retaining a reasonable weight, the panels are also reinforced, inside the fuselage, by reinforcing elements, in particular by frames 4, essentially positioned along sections of the fuselage substantially perpendicular to the longitudinal axis of the fuselage. Often, a reinforcing frame is placed on a face 32 of said ring opposite the face 31 in contact with the ends 22a, 22b of the panels 2a, 2b in order to reinforce the fuselage. Said frame is then fixedly attached to the ring 3 and to each panel 2a, 2b with the aid of the fasteners 7.
When the forces, having to be transmitted between the two panels of the fuselage by said ring, increase, in particular because of the increased dimensions and capacities of an existing aircraft, such as for example a lengthening of the fuselage, then the use of said ring to produce the assembly between said two panels is not satisfactory.
Specifically, the ratio between the fatigue dimensioning of the join and its static dimensioning, which determines a reserve factor, reduces when the loads in the join increase.
This reduction of the reserve factor results in the virtually inevitable and rapid appearance, in service, of fatigue cracks under alternating stress. This high risk of the rapid appearance of cracks will reduce the inspection interval. In addition, because of the complex architecture (number and stacking of parts) of the zone concerned, these cracks are very difficult to detect and the inspections will consequently be even more difficult to carry out, which may cause the first inspection to be brought forward, and then the inspection intervals to be shortened.
A known solution consists in increasing the thickness of the ring to increase the ability of the join to transmit the forces between the two panels. However, the thickening of the ring does not significantly and proportionally improve the reserve factor, the reserve factor still remaining low, in particular because of the not inconsiderable increase in the secondary bending moment in the join, a consequence of the thickening of the ring.
The level of forces that it is now likely to encounter in heavily loaded joins can therefore no longer be transmitted by an assembly of the type used in the prior art, as shown in FIG. 1, which works in a simple shearing manner.
The use of a new assembly of panels making it possible to withstand ever-increasing loads is important, while ensuring a satisfactory reserve factor.